1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a turbine rotor blade with enhanced leading edge impingement cooling.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine includes a turbine with multiple rows or stages of rotor blades that react with a high temperature gas flow to drive the engine or, in the case of an industrial gas turbine (IGT), drive an electric generator and produce electric power. It is well known that the efficiency of the engine can be increased by passing a higher temperature gas flow into the turbine. However, the turbine inlet temperature is limited to the material properties of the first stage vanes and blades and the amount of cooling that can be achieved for these airfoils.
In latter stages of the turbine, the gas flow temperature is lower and thus the airfoils do not require as much cooling flow. In future engines, especially IGT engines, the turbine inlet temperature will increase and result in the latter stage airfoils to be exposed to higher temperatures. To improve efficiency of the engine, low cooling flow airfoils are being studied that will use less cooling air while maintaining the metal temperature of the airfoils within acceptable limits. Also, as the TBC (thermal barrier coating) gets thicker, less cooling air is required to provide the same metal temperature as would be for a thicker TBC.
FIG. 1 shows an external pressure profile for a turbine rotor blade. As indicated in the figure, the forward region of the pressure side surface experiences high hot gas static pressure while the entire suction side of the airfoil is at a much lower hot gas static pressure than the pressure side. The pressure side pressure profile in the line on the top while the suction side pressure profile is the line on the bottom in the FIG. 1.
FIG. 2 shows a prior art turbine rotor blade with a (1+5+1) forward flowing serpentine cooling circuit for a first stage rotor blade. FIG. 3 shows a schematic view of the rotor blade of FIG. 2 and FIG. 4 shows a flow diagram of the flow path through the FIG. 2 rotor blade. The prior art blade cooling circuit includes a leading edge cooling supply channel 21 connected to a leading edge impingement cavity 23 by a row of metering and impingement holes 25, and where the impingement cavity 23 is connected to a showerhead arrangement of film cooling holes 26 and gills holes 24 on both sides to discharge a layer of film cooling air onto the leading edge surface of the airfoil. A forward flowing 5-pass serpentine cooling circuit is used in the airfoil mid-chord region with a first leg 11 for supplying cooling air located adjacent to the trailing edge region of the airfoil. The second leg 12, third leg 13, fourth leg 14 and fifth leg 15 of the serpentine flow toward the leading edge in series with rows of film cooling holes 17 connected to some of the 5 legs to discharge film cooling air onto the pressure or suction sides of the airfoil. A trailing edge cooling air supply channel 31 supplies cooling air for the trailing edge region and is connected to a series of impingement holes 32 and 34 to first and second impingement cavities 33 and 35, which is connected to a row of exit holes or slots 36 to discharge the spent impingement cooling air. Film cooling holes 37 can also be connected to the impingement cavity 33.
The cooling air flows from the trailing edge region toward the leading edge region and discharges into the hot gas side pressure section of the pressure side of the airfoil. In order to satisfy the back flow margin criteria, a high cooling supply pressure is needed for this particular design, and thus inducing a high leakage flow. In the prior art cooling arrangement of FIG. 2, the blade tip section is cooled with double tip turns in the serpentine circuit and with local film cooling. Cooling air bled off from the 5-pass serpentine flow circuit will thus reduce the cooling performance for the serpentine flow circuit. Independent cooling flow circuit is used to provide cooling circuits from the 5-pass serpentine flow circuit is used for cooling of the airfoil leading and trailing edges.
As the TBC technology improves and more industrial turbine blades are applied with thicker or low conductivity TBC, the amount of cooling flow required for the blade will be reduced. As a result, there is not sufficient cooling flow for the prior art design with the 1+5+1 forward flowing serpentine cooling circuits of FIG. 2. Cooling flow for the blade leading edge and trailing edge has to be combined with the mid-chord flow circuit to form a single 5-pass flow circuit. However, for a single forward flow 5-pass circuit with total blade cooling flow BFM (back flow margin) may become a design problem.